Technology development plan - NEOShield-2...WP 12 Deliverable D12.3 Technology development plan WP...

This project has received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 640351.

NEOShield-2 Science and Technology for Near-Earth Object Impact Prevention

Grant agreement no: 640351 Project Start: 1 March 2015

Project Coordinator Airbus Defence and Space DE Project Duration: 31 Months

WP 12 Deliverable D12.3

Technology development plan

WP Leader DLR Task Leader ADS-DE Due date M30, 31 Aug 2017 Delivery date 30.09.2017 Issue 1.0 Editor (authors) Kilian Engel Contributors NEOShield-2 Team Verified by Albert Falke Document Type R Dissemination Level PU

The NEOShield-2 Consortium consists of: Airbus Defence and Space GmbH (Project Coordinator) ADS-DE Germany Deutsches Zentrum für Luft- und Raumfahrt e.V. DLR Germany Airbus Defence and Space SAS ADS-FR France Airbus Defence and Space Ltd ADS-UK United Kingdom Centre National de la Recherche Scientifique CNRS France DEIMOS Space Sociedad Limitada Unipersonal DMS Spain Fraunhofer-Gesellschaft zur Förderung der angewandten Forschung e.V. EMI Germany GMV Aerospace and Defence SA Unipersonal GMV Spain Istituto Nazionale di Astrofisica INAF Italy Observatoire de Paris OBSPM France The Queen’s University of Belfast QUB United Kingdom

NEOShield-2 D12.3 Title: Technology development plan Issue, Date i1.0, 30.09.2017 /19

Change Record

Issue Date Section, Page Description of Change

0.9 21.09.2017 Draft version for consortium review.

1.0 30.09.2017 Incorporation of consortium internal review updates.


Table of Contents

1 Introduction ............................................................................................................................................. 4

1.1 Scope .................................................................................................................................................. 4

1.2 List of Abbreviations .................................................................................................................... 4

1.3 Applicable Documents ................................................................................................................ 5

1.4 Reference Documents ................................................................................................................. 5

2 Missions and required functionalities ............................................................................................ 6

3 Overview of Technologies and TRL ............................................................................................... 11

4 Recommendations for future developments ............................................................................. 16

Annex ................................................................................................................................................................. 19


1 Introduction

1.1 Scope The purpose of this technology roadmap is:

• to provide a compact overview of the technological capabilities required for a set of NEO-related reference missions

• assess the development status of these technical capabilities • identify needed activities to bring the identified technologies to TRL 6

The set of reference mission includes the missions studied under the NEOShield and NEOShield-2 projects, as well as other NEO missions that are deemed relevant to mitigation, characterisation, or utilisation. Missions are included for which sufficient information is available from work under NEOShield(-2) or other study activities of the NEOShield-2 consortium. The focus is on demonstration missions rather than operational missions (e.g. real threat mitigation), which in absence of a real threat are seen as the intermediate goal of most technology development activities.

The technology readiness level evaluation is based primarily on a European perspective, i.e. gives an assessment of the technological capabilities within Europe. This allows an assessment of the development needs that are implied by the decision to pursue different missions or mission contributions (e.g. in international collaborations) by European actors. The assessment focuses on the development needs in order to reach TRL 6, since this is the ideally required level to support an implementation decision with a reasonable amount of schedule and financial risk. Higher TRL generally implies the decision to implement the technology in an actual mission.

The document also gives some preliminary recommendations about how the technology development can be focused and prioritised, while pointing out that this is obviously also a decision based on programmatic / mission priorities.

1.2 List of Abbreviations AD Applicable document

AIDA Asteroid Impact and Deflection Assessment (Mission)

AOCS Attitude and orbit control system

CAM Collision avoidance manoeuvre

CoG Center of gravity

EM Engineering model

ESOC European Space Operations Center

FDIR Failure detection isolation and recovery

FM Flight model

FOV Field of view

GNC Guidance Navigation and Control

GSFC Goddard Space Flight Center

GT Gravity tractor

HIL Hardware in the loop

IBS Ion-beam shepherd

KI Kinetic Impactor


LEO Low earth orbit

MIL Model in the loop

NEO Near Earth Object

NEOTωIST Near-Earth Object Transfer of angular momentum (ω) Spin Test

OBC Onboard computer

ODM Orbital departure manoeuvre

PIL Processor in the loop

QM Qualification model

RD Reference document

S/C Spacecraft

SIL Software in the loop

TRL Technology readiness level

1.3 Applicable Documents [AD1] NEOShield-2: “Science and Technology for Near-Earth Object Impact Prevention”, Grant

Agreement no. 640351, 28.10.2014.

1.4 Reference Documents [RD1] NEOShield-2, D3.4 GNC Technology Validation Test Report, Issue 1.1, 21.09.2017

[RD2] NEOShield-2, D8.2 Report on NEO Orbit Determination and Monitoring Strategy, Algorithms Design and Functional Validation, Issue 1.0, 30.06.2017

[RD3] Memo: Examination of asteroid flyby visual navigation - feasibility and performance, 31.08.2017, DeMichele Emmanuel

[RD4] NEOShield-2, D6.1 Technologies for Return and In‐situ Analysis of NEO Samples, Issue 1.0, 26.01.2016

[RD5] NEOShield-2, D6.3 Sampling Device: Design and Test Report, Issue 1.1, 29.04.2016

[RD6] NEOShield D8.2 NEOShield Kinetic Impactor Demonstration Mission Detailed Design, Issue 1, January 2015

[RD7] Asteroid Impact & Deflection Assessment mission ; ESA website ; http://www.esa.int/Our_Activities/Space_Engineering_Technology/Asteroid_Impact_Mission/Asteroid_Impact_Deflection_Assessment_mission ; retrieved Sept. 2017

[RD8] Engel et. al., NEOTωIST - An Asteroid Impactor Mission Featuring Sub-spacecraft for Enhanced Mission Capability, IAC-16,B4,8,7,x34163, International Astronautical Congress 2016

[RD9] NEOShield, D8.4 NEOShield Blast Deflection Demonstration Mission Detailed Design, issue 2.0, 20 February 2015

[RD10] NEOShield, D7.1: Gravity Tractor Feasibility Report, issue 1.0, 1 March 2013

[RD11] NEOShield, D7.5.4: Ion Beam Shepherd Deflection Concept, issue 1.0, 15 Sept 2013

[RD12] NEOShield-2, Reference Mission Definition: Sample Return, issue 2.0, 23.06.2016


2 Missions and required functionalities This chapter provides a short description of the reference mission concepts that are considered. At the end there is a summary of the key functionalities/ technical capabilities implied by each mission concept.

(A) Classic Kinetic Impactor demonstration mission - classic

The space segment of this mission consists of two spacecraft, the Explorer and the Impactor spacecraft. Depending on the chosen design, launch and cruise of both spacecraft may be designed to be joint or separate. The Explorer arrives and rendezvous with the NEO prior to the Impactor and performs a physical of the object characterisation (mass, CoG position, rotational state, topography and surface properties) before the Impactor arrives. This includes in particular a precision orbit determination of the NEO, as well as observation of the object geometry and size. The latter is used for improved Impactor targeting accuracy, using visual navigation. After completion of this NEO characterisation phase the Impactor spacecraft arrives and impacts the NEO at hypervelocity, thus transferring momentum to the target NEO and changing its orbit slightly. In many cases the impactor will consist of the mission spacecraft as well as the transfer or launcher upper stage, in order to increase the impacting mass. The impact itself will be observed by the Explorer from a safe position. After the impact, the Explorer performs another precision orbit determination in which together with the NEO properties allows quantification of the momentum change imparted by the Kinetic Impactor.

For details of the NEOShield KI mission see [RD6]

(B) Classic Kinetic Impactor demonstration mission - classic with surface science

The two-spacecraft KI mission architecture described above, can also be expanded and enhanced by a small mobile surface science probe. In this case the Explorer spacecraft would deploy a small probe for in-situ surface characterisation of the target object, e.g. for improved understanding of NEO composition and structure. This surface science probe may be designed for a passive or (semi-)controlled touchdown. The later significantly reduces the demands on the Explorer spacecraft, since the small probe can be deployed more safely from a greater distance. The Asteroid Impact and Deflection Assessment Mission (AIDA), as originally foreseen, is an example of this mission type [RD7].

(C) Kinetic Impactor Demonstration Mission - NEOTωIST concept

NEOTωIST stands for Near-Earth Object Transfer of angular momentum (ω) Spin Test. This describes a demonstration mission intended to develop the capabilities required to execute an effective kinetic impactor NEO deflection mission. The chosen measurement technique and employment of small sub-spacecraft for observation purposes represent a novel approach to achieving the main goals of such a demonstration mission. The approach promises comparatively low cost and features capabilities that are unique and valuable for an operational deflection mission. Most standard deflection demonstration missions propose to quantify momentum transfer from the impactor spacecraft to the target object by measuring a change in its heliocentric orbit. The change is typically so small that it must be performed via radio-science from a second observer spacecraft which rendezvous with the NEO prior to impact. In our case the NEO is struck off-center which changes its spin rate. This rate change, which can be measured from Earth via light curve measurements, allows quantification of the transferred momentum. Using this measurement method the need for an observer spacecraft for the purpose of NEO orbit measurement is eliminated. The second function of the observer spacecraft is the close-up observation of the impact event for improvement of impact effectiveness modelling. The NEOTωIST mission achieves this observation by deploying several


small sub-spacecraft from the main impactor spacecraft shortly before impact. These sub-spacecraft allow observation of the impact event from multiple vantage points some of which are unique because their destruction is accepted. At least one sub-spacecraft trajectory is planned such that survival is guaranteed, which enables it to receive observation data from the other spacecraft for delayed transmission to Earth.

For details on the NEOTωIST mission concept see [RD8].

(D) Blast deflection (with & w/o surface science)

Blast deflection missions are a category of missions in which a blast in most is induced on or near the surface of an NEO. The NEO material which is either blasted off the surface or vaporized carries momentum thus imparting momentum on the NEO and changing its orbit. In many cases the mechanism of inducing a sufficiently powerful blast is through a nuclear device. Typically, the mission architecture would consist of an Explorer spacecraft which characterizes the NEO and determines its orbit precisely before and after the blast event to verify and quantify the deflection. In this case the blast device is carried by a sub-spacecraft which is deployed near the NEO surface by the Explorer. While the sub-spacecraft positions and triggers the blast device, the Explorer attains a safe position. Details on the blast deflection concept studied under NEOShield can be found in [RD9].

(E) Orbiting Explorer mission / Remote prospector

This mission scenario describes that of a spacecraft which performs a rendezvous and observes and characterizes the NEO remotely using a suite of remote sensing instruments. This would typically also include precision orbit determination and gravity measurements via trajectory measurements. This type of Explorer mission will in many cases be part of other missions such the KI missions described above. A stand-alone example is the Dawn mission. The mission scenario and capabilities are also representative of an orbital prospector mission in the context of resource prospecting/extraction scenarios. The instrumentation will obviously be tailored to the priorities of the mission.

(F) Gravity tractor

This “slow-push” concept is a technique for fine-tuning the orbit of an asteroid that has already been deflected by other means. It exploits one or more solar-electric propulsion engines to provide a gradual thrust that changes the NEO target and the S/C orbits together, maintaining a fixed distance between the two elements during the whole operations performance. The force from the spacecraft is applied to NEO via gravitational attraction. To achieve a reasonable force this implies a certain minimum spacecraft mass. The thrust must be applied along or against the orbital vector in order to “affect” the orbital parameters of the NEO target, and in particular to change its period. However, the thrusters must be canted to prevent propellant from hitting the asteroid (which would push it away). The longer the thrust is applied, the greater the change in period will be. And the longer the NEO target is observed, the more apparent this period change will be. The GT in practice can simultaneously make small adjustments to the asteroid orbit and provide the information needed for a very precise tracking of it from the Earth. In other words, it almost matches the functionalities of a KI observer or explorer S/C as described in the previous sections.

Taking into account the above considerations, it turns out that a Gravity Tractor mission is expected to be designed and launched for the following purposes:

• to refine the asteroid deflection achieved by using another deflection spacecraft, enabling a better and more controlled tuning of the NEO target orbital parameters

• to support the operations of another deflection spacecraft, e.g. the Kinetic Impactor, providing advanced GNC and monitoring capabilities, generating data about the asteroid


orbital parameters before and after the use of another deflection technique and also during the use of the Gravity Tractor concept.

Details on the concept studied under NEOShield can be found in [RD10].

(G) Ion-beam shepherd

The Ion-Beam Shepherd is able to gradually deflect an asteroid, simply by firing propellant at the asteroid, to produce a gradual change in the asteroid’s momentum and velocity. A counteracting thrust is required to prevent the Ion-Beam Shepherd from drifting away. Both the thrust and the beam hitting the asteroid would typically be created by an electric thruster. Like the gravity tractor method, the Ion-beam shepherd allows fine tuning the asteroid trajectory very precisely, and has low sensitivity to the cohesive properties or rotation of the target object.

For details on the concept see [RD11].

(H) In-situ Explorer mission / In-situ prospector

This scenario describes a mission which puts a vehicle on the surface of a NEO which carries a suite of sensors and laboratory instrumentation for in-situ analysis of the object properties. The architecture may be either a single vehicle which performs the cruise, rendezvous and surface operations, or in many cases a two vehicle design with an orbiter and a mobile surface probe deployed from the orbiter. In this case some mission designs conceivably allow for a relatively small surface probe which has the potential to simplify the mission and reduce overall cost. In both architectures the fundamental functionalities are similar regardless of whether they are provided by a single vehicle or distributed among two. In most cases this mission scenario will also perform orbital remote characterisation for a synoptic view. The surface instrumentation may include for instance the, drills, mass spectrometers, cameras, microscopes, thermal sensors.

The mission scenario is broadly representative of missions with several different types of focus, i.e. such surface examinations may serve deflection preparation, scientific interests or resource prospecting purposes.

This mission scenario has not been explicitly studied in detail under the NEOShield(-2) projects.

(I) Sample Return mission

This mission describes a scenario where a sample of Asteroid material is extracted from the target object and returned to Earth for laboratory examination. The typical architecture consists of a single vehicle which rendezvous with the Asteroid, performs some orbital characterisation, touches down for sample extraction, then returns to Earth and releases the sample in an entry capsule for atmospheric entry descent and landing. Finally the capsule and sample is retrieved on ground. Depending on the objectives and design the sample extraction may be from the surface or subsurface of the NEO. Further, different extraction methods either require full landing and standstill on the surface or allow for a "touch and go" approach where the sample is extracted during a brief non-stationary surface contact. The later method is the one employed by the most prominent mission examples already implemented, that is Hayabusa and Osiris-Rex. The NEOShield-2 work on a sample return mission scenario is summarized in [RD12].

(J) Flyby Characterisation / Mini remote prospector

This mission describes a scenario where a spacecraft encounters a NEO on a flyby trajectory. The relative velocity is typically several km/s. In cases this is the fastest and cheapest (low delta-v, small vehicle) possibility to gain some preliminary information about an unknown object. This may include improvement of the NEO trajectory, constraining the size, mass, geometry and possible composition. The mission scenario is typically discussed when information about a


potentially hazardous object is required quickly and a rendezvous trajectory would take too long to achieve or the more complex mission is not justified by the impact probability. Such missions may also have a role to play for preliminary resource prospecting even though the value to cost is yet to be determined. The mission utility is in some cases significantly lower than for a rendezvous mission. Consequently it is attractive to consider small/ low cost vehicle designs, especially in scenarios where no imminent threat is present. Such a vehicle would share many features with the Flyby Module proposed under the NEOTωIST scenario.

Table 2-1: Major mission functions by mission type

A B C D E F G H I J

# Function / Mission KI

Dem

o cl

assi

c

KI D

emo

with

surf

ace

scie

nce

KI D

emo

NEO

TωIS

T

Blas

t def

lect

ion

(with

&

w/o

surf

ace

scie

nce)

Orbi

ting

Expl

orer

Grav

ity tr

acto

r

Ion

beam

shep

herd

In-s

itu E

xplo

rer

Sam

ple

Retu

rn

Flyb

y ch

arac

teri

zatio

n

Guidance, Navigation and Control

1 Impactor GNC x x x

2 Reconnaissance/ Orbiter GNC 1 x x x x x x x x

3 Autonomous proximity hovering GNC8 x x

4 Landing / Sample Return GNC x (x)2 x x

5 Precision NEO orbit determination x x x x x x (x)3 (x)3

6 Flyby navigation x x

Observation and analysis capabilities

7 Orbital observations - ejecta (x)4 (x)4 (x)4

8 Orbital observations - crater x x x

9 Orbital observations - NEO physical properties x x x x x x x x

10 Flyby observations -ejecta x

11 Flyby observations - NEO physical properties x x

12 Surface science package x (x)2 x

13 Earth-based determination of rotation period x

Various platform and mechanism capabilities

14 Small deep space platform x x (x)2 (x)5 X6

15 Surface mobility / multisite capability (x)7 (x)7

16 Sample extraction and storage x


A B C D E F G H I J

17 Ion-propulsion / Ion-shepherding beam x x

18 Blast deflection payload x

1) Reconnaissance orbiter GNC contains the sub-functionalities of: Close Approach, Inertial Hovering, Body-fixed hovering (See [RD1] for details)

2) If blast application is preceded by surface reconnaissance

3) Nice to have

4) Ejecta observation may not be possible safely for orbiting S/C, depends on mission concept, e.g. sub-spacecraft could be employed

5) The assumed mission concept would be more affordable with a small mobile surface probe

6) The mission utility of a Flyby characterisation is comparatively lower, such that a "small" low-cost platform enables more reasonable mission utility to cost ratio

7) Surface mobility/ multi-site capability would potentially add significant value, but is optional depending on the scope of the mission and the precise questions to be addressed

8) The GNC required for slow push methods (GT and IBS) is considered different and more challenging than the typical orbiter GNC, since the spacecraft must remain stationary much closer to the neo Surface (< 100s of m) for much longer periods (months - years)

________________________


3 Overview of Technologies and TRL This section provides an assessment and justification of the current TRL for the individual technical capabilities that are given outlined in the previous chapter.

Table 3-1: Evaluation technical capabilities by TRL status and needed development effort

# Mission / Function TRL Justification of TRL Main points needed to reach TRL 6

/other study needs Dev. effort

Guidance, Navigation and Control

1 Impactor GNC 5

MIL, PIL, HIL test campaigns. Details in D3.4 [RD1]. Closed-loop simulation missing for to reach TRL 6.

- Closed loop simulation - Use of flight representative camera hardware Medium

2

Reconn-aissance/ Orbiter GNC

5-6

MIL, PIL, HIL test campaigns. Details in D3.4 [RD1].

- To mitigate the scaling effects the following points shall be taken into consideration:

1)Make use of existing/available narrow angle camera hardware for Close Approach and Arrival Inertial Hovering for 6h phases (not the case in Body-Fixed Hovering phase using wide angle camera)

2) Include real altimeter in the verification chain (considering ways to overcome the limitations in available HIL test benches)

- Perform full verification of the scenarios including FDIR functions where CAM manoeuvres shall be triggered in the HIL campaign.

- The software testbed shall also compatible in form and fit with the FM/EM OBC.

- The costs to realize the QM/FM models shall be fully determined.

(See [RD1])

Medium

3

Autonomous proximity hovering GNC

2-3

Not studied in NEOShield-2, no significant capability known in Europe

Full development effort required including. MIL, PIL, HIL campaigns, verifying critical functions in relevant environment and with flight representative hardware

Signific.




4

Landing / Sample Return GNC

5-6

MIL, SIL, PIL, HIL test campaigns. Details in D3.4 [RD1]

The following recommendations have been worked out to increase the achieved TRL to 6.

-To mitigate the scaling effects the following points shall be taken into consideration:

1) The right FOV shall be taken for the final validation.

2) Dedicated hardware tests are proposed to have a validated altimeter model.

3) The landing conditions shall be harmonized.

-CAM manoeuvres shall be triggered in the HIL campaign.

-The software testbed shall also compatible in form and fit with the FM/EM OBC.

-The costs to realize the QM/FM models shall be fully determined.

Software tested on breadboard processor unit which is representative in terms of function. Not representative in terms of form and fit. No radiation hardened components have been used.

(See [RD1])

Medium




5

Precision NEO orbit deter-mination

5

-Development of on-ground processing algorithms to perform orbit determination and monitoring of small NEOs with very weak gravity fields. - The technology development has been validated up till a technology readiness level of TRL5 achieved by validation/demonstration of the technology in a relevant environment. In a first step, the validation of the relative referencing link was performed using HIL test facility data. Thereafter, the validation was performed using simulated mission data covering cases not addressed in the HIL test facility. ODM on-ground and momentum enhancement factor estimation algorithms are implemented in a Matlab & Simulink language. Details in D8.2 [RD2].

A more profound feasibility study involving the ESOC and/or GSFC teams to further refine the interface definition. Porting them to the operational environment at ESOC and/or GSFC would allow reaching TRL6.

Medium

6 Fly-by navigation 3

Setup of filter and simulation that shows basic feasibility. (See [RD3])

- Confirmation of hardware assumptions, in particular camera. - Modelling of AOCS performance and contribution to navigation errors. -Full MIL, PIL, HIL test campaigns.

Signific.

Observation capabilities

7

Orbital obser-vations - ejecta

3

It is assumed that ejecta can be observed by means of optical camera. Deep impact shows that this is fundamentally possible. However, the understanding of the optical properties of the expected NEO ejecta is not yet sufficiently well understood to judge the performance of different available cameras.

- Develop reliable models of optical properties of ejecta. -Check these against available camera hardware for feasibility. - Verify performance for selected hardware. Signific




8

Orbital obser-vations - crater

5

General performance requirements consistent with existing optical cameras. However, no actual equipment selected.

Refinement of performance requirements. Selection, performance verification, and qualification of actual camera system.

Modest

9

Orbital obser-vations - NEO physical properties

(4)

Evaluation depends on precise observables to be assessed. Heritage exists for a good portfolio of remote sensing equipment.

- Refine selection of observables and performance requirements. - Selection, performance verification, and qualification of actual instrument hardware.

Signific

10

Flyby obser-vations -ejecta

3

It is assumed that ejecta can be observed by means of optical camera. Deep impact shows that this is fundamentally possible. However, the understanding of the optical properties of the expected NEO ejecta is not yet sufficiently well understood to judge the performance of different available cameras.

- Develop reliable models of optical properties of ejecta -Check these against available camera hardware for feasibility - Verify performance for selected camera hardware - Development of imaging pointing mirror mechanism to EM level

Signific

11

Flyby obser-vations - NEO physical properties

4

Evaluation depends on precise observables to be assessed. Heritage exists for a good portfolio of remote sensing equipment.

-Refine selection of observables and performance requirements. -Selection, performance verification, and qualification of actual instrument hardware

Signific

12 Surface science package

(4)

Evaluation depends on precise instrumentation and desired analysis. Heritage exists for a good portfolio of remote analysis equipment.

- Refine selection of observables and performance requirements. - Selection, performance verification, and qualification of actual instrument hardware. - For small surface science probe concepts, downscaling of some existing instrument concepts may be required.

Signific

13

Earth-based determi-nation of rotation period

9

Light curve measurements for rotation period determination have been performed with sufficient accuracy.

n/a

Low




Various platform and mechanism capabilities

14 Small deep space platform

4

Flexible LEO Platform (FLP2) development activities at Airbus. High performance small LEO platform. Software verification facility. Hardware partially integrated.

Design and component updates for mission specifics and radiation hardness. Software modification for deep space applications. Functional testing of new functionalities, environmental testing, especially radiation. Deep space communications hardware. Selection or development of suitable propulsion system (chemical or electrical).

Significant

15

Surface mobility / multisite capability

3

Early trade studies and conceptual designs for surface mobility concepts (e.g. CORE concept, Astrone).

-Selection and refinement of vehicle concept. -Selection of thruster technology and configuration. -Development and verification of GNC concept.

High

16

Sample extraction and storage

3-7

Various concepts, in various stages of development. See [RD4] for details. NEOShield development at TRL 4-5 [RD5]

Selection of specific sampling principle based on mission objectives and constraints. Development program to TRL 6

Significant-High (depend

s on type).

17

Ion-propulsion / Ion-shepher-ding beam

7

Assumption is that QintiQ T6 would be a good candidate for a demonstration mission. NEOShield looked at a scaled up version of the T6 for an operational mission. The T6 is under qualification for the BepiColombo program

-No technology development anticipated. -Detailed study to confirm preliminary assessment that T6 is suitable for a demonstration mission. Low

18 Blast deflection payload

3

No heritage for blast deflection payloads in Europe. Some conceptual work under NEOShield Project. Some GNC and platform functionalities can be reused for deployable free flying payload.

- Selection of blast principle. - Development of key technologies to TRL 6

High


4 Recommendations for future developments Development priorities for technologies clearly depend on desired missions and capabilities, which is ultimately a political and programmatic decision which must be made by the stakeholders in Europe, likely considering the positions and plans of international partners.

Absent of a clear prioritisation of specific missions we can order the above described capabilities into three broad categories which also represent a tentative prioritisation.

Category 1:

These are capabilities that are applicable to many (almost all missions), which makes them inherently valuable to a NEO exploration and protection programme, independently of the precise mission priorities. As visible in Table 2-1, this includes the following:

• #2 Orbiter GNC • #5 Precision NEO orbit determination • #8 Orbital observations - crater (in many cases included in #8) • #9 Orbital observations of NEO physical properties • #13 Earth-based determination of rotation period (included because valuable science

tool and already quite mature) • #14 Small deep space platform

In most cases the remaining effort to TRL 6 is also limited. The wide applicability and generally modest development effort makes these capabilities potential development priorities as long as specific missions have not been prioritised.

Category 2:

The capabilities in this category are generally more mission specific, and on average need more development effort but also offer high value in terms of knowledge return. The capabilities are largely those associated with the classic KI scenario, the NEOTωIST scenario or the in-situ/sample return study of asteroid material. This includes the following capabilities:

• #1 Impactor GNC • #3 Autonomous proximity hovering GNC • #4 Landing GNC • #6 Flyby navigation • #7 Orbital observations-ejecta • #10 Flyby observations-ejecta • #11 Flyby observations- NEO physical properties • #12 Surface science package • #15 Surface mobility / multi-site capability • #15 Sample extraction and storage

Note that #15, #4, #12 implicitly describe a small surface science probe which is a valuable yet relatively affordable asset in many mission scenarios.

Category 3:

This summarizes technologies which can be considered low priority for the European NEO effort, for one of the following reasons:

• Already being developed in the context of other activities anyway #17 Ion propulsion/ shepherding beam

• Currently judged to be low priority activities in Europe (# 18 Blast deflection payload)


Recommendations

The following recommendations are made with regard to technology development:

1) Define at a political and programmatic level, with scientific and engineering input, specific European reference mission(s) or mission contributions as soon as possible. Table 4-1 shows a preliminary categorization of mission scenarios according to main objectives and relative funding required. The NEOShield-2 consortium believes that a deflection demonstration mission is the critical next step in significantly moving ahead the state of preparation for a real threat scenario, i.e. a substantial increase in the deflection capability. Assuming that funding challenges for such a mission are non-trivial, the NEOTωIST concept is an attractive mission to study in depth since it promises somewhat lower cost than most other deflection demonstration missions. Alternatively or in addition, a mobile surface probe is a mission element that is an attractive contribution to surface science activities at asteroids and many other mission scenarios.

2) Revise prioritization of the associated capabilities based on selected mission(s). Absent of such definition pursue capabilities roughly in the sequence of Category 1-3

3) More specifically, it is recommended to initiate, as soon as possible, development programs to achieve TRL 6 for all Category 1 capabilities. This is recommended since it can be achieved with modest resources and is useful in most future programme scenarios.

4) It is recommended to initiate in depth system studies for at least two mission scenarios / system concepts. As discussed above NEOTωIST and a surface science probe would be attractive options.

5) It is recommended to simultaneously start development programmes for the technologies associated with these missions, with the goal of reaching TRL 6.

Note that some of the presented recommendations would need to be modified if the programmatic status of the AIDA programme changes (currently no European participation foreseen). Some of the recommendations would already be implemented in this case, i.e. to select a European mission contribution and develop the associated technologies.

Table 4-1: Most attractive reference missions depending on main interest and funding levels

Main interest / funding €€€ €€ €

KI deflection demonstration Classic KI Demo mission (optionally with surface science)

NEOTωIST

Scientific characterization / prospecting

Sample return Orbiter with small mobile surface probe

Small surface probe as mission contribution

Slow push deflection demonstration

Gravity tractor or IBS

Slow push deflection demonstration and scientific characterization

Gravity tractor or IBS with small mobile surface probe

Small surface probe as mission contribution

Table 4-2 outlines a rough timeline into which technology development tasks and system/ mission study tasks could be organized to support a sustainable and dynamic NEO deflection demonstration mission programme.


Table 4-2: Preliminary high-level timeline for technology development and mission study activities

Preliminar technology development timeline

Activity Start End YearQarter 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4

Development of Category 1 technology to TRL 6 KO (T0) KO + 2 years

Mission/ System Phase A/B1 study for NEOTωIST KO (T0) KO + 2 years

Mission/ System Phase A/B1 study for surface probe KO (T0) KO + 2 years

Development key Cat-2 technologies for NEOTωIST to TRL 5/6 KO (T0) KO + 3 years

Development key Cat-2 technologies for surface probe to TRL 5/6 KO (T0) KO + 3 years

Consolidation of selected mission concept K0 + 2 years KO + 3 years

Consolidation of technologies of selected mission K0 + 2 years KO + 3 years

Implementation of demonstration mission K0 + 3 years KO + 7 years

KO Programme

Mission selection

Start mission implement.

Mission launch

7 861 2 3 4 5


Annex

D12.3b NEOTωIST Mission Scenario and System Design Concept

This project has received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 640351.

NEOShield-2 Science and Technology for Near-Earth Object Impact Prevention

Grant agreement no: 640351 Project Start: 1 March 2015

Project Coordinator Airbus Defence and Space DE Project Duration: 31 Months

WP 12 Deliverable D12.3b

NEOTωIST Mission Scenario and System Design Concept

WP Leader DLR Task Leader ADS-DE Due date M30, 31 Aug 2017 Delivery date 30.09.2017 Issue 1.0 Editor (authors) Kilian Engel Contributors NEOShield-2 Team Verified by Albert Falke Document Type R Dissemination Level PU

The NEOShield-2 Consortium consists of: Airbus Defence and Space GmbH (Project Coordinator) ADS-DE Germany Deutsches Zentrum für Luft- und Raumfahrt e.V. DLR Germany Airbus Defence and Space SAS ADS-FR France Airbus Defence and Space Ltd ADS-UK United Kingdom Centre National de la Recherche Scientifique CNRS France DEIMOS Space Sociedad Limitada Unipersonal DMS Spain Fraunhofer-Gesellschaft zur Förderung der angewandten Forschung e.V. EMI Germany GMV Aerospace and Defence SA Unipersonal GMV Spain Istituto Nazionale di Astrofisica INAF Italy Observatoire de Paris OBSPM France The Queen’s University of Belfast QUB United Kingdom

NEOShield-2 D12.3b Title: NEOTωIST Mission Scenario and System Design

Concept Issue, Date i1.0, 30.09.2017 Page 2/33

Change Record

Issue Date Section, Page Description of Change

1.0 30.09.2017 All Issue of document



Table of Contents

1 Introduction ............................................................................................................................................. 4

1.1 Scope .................................................................................................................................................. 4

1.2 List of Abbreviations .................................................................................................................... 4

1.3 Applicable Documents ................................................................................................................ 5

1.4 Reference Documents ................................................................................................................. 5

1.5 Introduction of mission concept ............................................................................................. 5

2 Mission and spacecraft design ........................................................................................................... 6

2.1 To-level mission objectives and constraint ......................................................................... 6

2.2 NEOTωIST measurement concept and mission architecture ....................................... 7

2.3 Sub-spacecraft deployment and observation strategy .................................................... 9

2.3.1 Flyby geometry ........................................................................................................................................ 9

2.3.2 Tracking strategy .................................................................................................................................. 14

2.3.3 Operations timeline ............................................................................................................................. 15

2.4 Analysis of flyby navigation ................................................................................................... 16

2.4.1 Position estimation (w/o camera noise) .................................................................................... 16

2.4.2 Position estimation (with noise) .................................................................................................... 19

2.5 Communications concept ........................................................................................................ 22

2.6 Design of Impactor and spacecraft stack ........................................................................... 23

2.7 Design of Flyby Module ............................................................................................................ 24

2.7.1 7.1 Design drivers for FBM ............................................................................................................... 24

2.7.2 7.2 Flyby module payload ................................................................................................................. 25

2.7.3 7.3 Outline FBM platform and operations .................................................................................. 26

2.8 Design of Chaser spacecraft ................................................................................................... 28

2.8.1 8.1 Design drivers for Chaser .......................................................................................................... 28

2.8.2 8.4 Outline Chaser design ................................................................................................................. 29

3 Impactor Guidance Navigation and Control .............................................................................. 30

4 Observation and measurements .................................................................................................... 31

4.1 Target Asteroid ........................................................................................................................... 31

4.2 Observation principle and analysis .................................................................................... 31

4.3 Observability of momentum enhancement factor (beta) ........................................... 31

4.4 Observing the ejecta in the visual ........................................................................................ 31

4.5 Observing the impact in the infrared ................................................................................. 31

5 Baseline mission trajectories .......................................................................................................... 33



1 Introduction

1.1 Scope This document reports the results of the work that the NEOShield-2 Consortium performed on the NEOTωIST mission concept.

The focus of the work was in the following four areas:

• Analysis and design of spacecraft and mission operations • Impactor GNC analysis • Analysis of observation needs and observation concept • Mission and orbit analysis

1.2 List of Abbreviations AD Applicable Document

AU Astronomical Unit

C Constraint

CH Chaser

CMG Control moment gyros

DQ Don Quijote mission

FBM Flyby Module

FOV Field of view

GNC Guidance Navigation and Control

IM Impactor

KI Kinetic Impactor

LEO Low Earth Orbit

LPF Lisa Pathfinder mission

MGA Medium gain antenna

MO Mission Option

NEO Near Earth Object

NEOTωIST Near-Earth Object Transfer of angular momentum (ω∙I) Spin Test

O Objective

OBC Onboard computer

PSD Power Spectral Density

RD Reference Document

STR Star tracker

Tx transmit

w/o without



1.3 Applicable Documents [AD1] NEOShield-2: “Science and Technology for Near-Earth Object Impact Prevention”, Grant

Agreement no. 640351, 28.10.2014.

1.4 Reference Documents [RD1] NEOShield-2, D3.4 GNC Technology Validation Test Report, Issue 1.1, 21.09.2017

[RD2] NEOShield-2, D4.1 GNC Design Report for Impactor Mission, issue 2.2, 29.09.2016

[RD3] Drube et al. The NEOTωIST mission (Near-Earth Object Transfer of angular momentum spin test), Acta Astronautica127, 2016

[RD4] Eggl et al., NEOT⍵IST: Determining the Momentum Enhancement, IAA Planetary Defense Conference 2017

1.5 Introduction of mission concept NEOTωIST stands for Near-Earth Object Transfer of angular momentum (ω∙I) Spin Test, and is a concept for a kinetic impactor demonstration mission, which aims to change the spin rate of an asteroid by impacting it off-center [RD3]. The change would be measured by means of lightcurve measurements with Earth-based telescopes. In contrast to most other kinetic impactor demonstration mission concepts, NEOTωIST does not require a reconnaissance spacecraft to rendezvous with the target asteroid for orbit change and impact-effect measurements, and is therefore a relatively inexpensive alternative.

The NEOTωIST mission would determine the efficiency of momentum transfer (the β-factor) during an impact, and help mature the technology required for a kinetic impactor mission, both of which are important precursor measures for a future space mission to deflect an asteroid by collisional means in an emergency impact hazard situation.

(25143) Itokawa has been chosen as the target for a mission study performed by the NEOShield and NEOShield-2 Consortia in 2014-2017.

The mission concept features several novel aspects. The most important of these are a new technique to quantify momentum transfer from the impacting spacecraft to the NEO, and the option to use small sub-spacecraft for observation purposes. The demonstration mission can be performed with just a spacecraft impacting the asteroid off-center (the Impactor) and the result being measured from ground. To gain additional information about the impact geometry and impact physics, the NEOTωIST Impactor can be made to deploy one or several sub-spacecraft from the main Impactor spacecraft shortly before impact. These sub-spacecraft allow observation of the impact event from multiple vantage points some of which are unique because their destruction is accepted.

The mission can be designed in several ways depending on the budget:

1. An impactor spacecraft only. The Impactor will cause the transfer of momentum to the asteroid and be responsible for close-up observation of the impact point before the impact.

2. A two-part spacecraft, which separate before impact into an Impactor and a Flyby module. The Flyby will observe the impact event and study the ejecta cloud and act as a data relay for the Impactor.

3. The two-part spacecraft with additional small sub-spacecraft, the Chasers, being ejected from the Impactor. The Chasers will observe the impact and the ejecta cloud from inside the plume and if there is a clear optical path to the impact point, then also the crater.

Overall, the concept promises comparatively low cost and features capabilities that are unique and valuable for an operational deflection or reconnaissance mission.



2 Mission and spacecraft design

2.1 To-level mission objectives and constraint Although there is little doubt about the principle feasibility of NEO deflection using a Kinetic Impactor, preparation for a real deflection mission requires the development and validation of the associated technology. Significant functionalities to be developed include the GNC to hit the target accurately, as well as means to verify an effective impact. Both are addressed by the proposed NEOTωIST mission concept.

Further, the models used to predict the impact dynamics need to be improved and validated. Currently, the observational data needed to do this is not available for hypervelocity impacts of relevant size. The relevance of accurate impact modelling is that it allows quantification of the momentum change imparted on the target object, i.e. deflection effectiveness. This in turn is needed to define spacecraft mass, geometry, and impact velocity such that sufficient alteration of the NEO trajectory is ensured. The uncertainty in the amount of moment transfer achieved stems primarily from different hypotheses about the momentum of the material ejected at impact, which typically augments the momentum transfer to the target object. The amount of additional imparted momentum achieved by the ejecta with respect to the momentum of the Kinetic Impactor is characterised by the beta factor

where pKI is the momentum carried by the Impactor and pEj is the momentum of the ejecta in opposite direction of the impact velocity. For the reasons explained, prediction of the β-factor is an essential parameter for deflection mission design. The NEOTωIST mission concept is conceived to measure the β-factor and to perform additional observations which allow to validate and constrain models of the impact dynamics.

Next to the technical challenges of the mission, the mission concept must respect programmatic constraints. The likelihood of implementation will depend on cost and implementation flexibility.

The top-level mission objectives and constraints of the proposed mission can be summarized as shown in the tables below.

Table 2-1: Major NEOTωIST Mission Objectives

# Objective Derived functionality

O1 Technology demonstration Kinetic Impactor

Impact target NEO with a spacecraft in hypervelocity regime with sufficient accuracy to ensure momentum transfer

O2 Technology demonstration of an observer spacecraft for impact verification

Demonstrate observation, from a flyby vehicle, of the impact event with sufficient quality to verify that the impact took place as required for deflection

O3 Deflection validation Measure target NEO orbit or rotation before and after impact to prove transfer of (angular) momentum.

O4 β determinat. / quantification of momentum transfer augmentation from ejecta

Quantify the magnitude of momentum carried by the ejecta

𝛽𝛽 = 𝑝𝑝𝐾𝐾𝐾𝐾 + 𝑝𝑝𝐸𝐸𝐸𝐸

𝑝𝑝𝐾𝐾𝐾𝐾



# Objective Derived functionality

O5 Observational data to validate/ improve impact modelling

Measure the dynamics and effects of the impact event Note: Observables and accuracy is subject to selection based on utility/cost assessment.

Table 2-2: NEOTωIST Mission Constraints

# Constraint Consequence

C1 Mission cost

Mission design and launcher selection shall be compatible with reduced budget compared to conventional deflection demonstration mission

C2 Cost & partnering flexibility

Mission concept shall allow tailoring according to budget, and offer opportunity for partnering

C3 Flexibility of implement. timeline

Multiple mission opportunities shall exist over an extended period of time to be compatible with different programme timelines

2.2 NEOTωIST measurement concept and mission architecture The mission concept is largely driven by Objectives 4 and 5 (see previous section), with the other objectives implicitly satisfied by the proposed concept.

Achievement of objective 4 means determination of the magnitude of the ejecta momentum vector caused by the KI impact. Standard deflection demonstration missions quantify momentum transfer from the impactor spacecraft to the target object by measuring a change in its heliocentric orbit. In the case of the proposed NEOTωIST concept the NEO is struck off-center which changes its spin rate (illustration in Figure 2-1:). The induced spin rate change can be measured accurately by Earth-based telescopes via light curve measurements, constituting a highly robust retrieval method. This is further described in [1].

Figure 2-1: Impact geometry and target object (Itokawa)

Relating the observed spin rate change to the momentum of the ejecta requires additional information about the mass properties of the object and the impact event geometry. The



observation concept and mission architecture are largely defined by the need to glean this information.

The information about NEO properties required for targeting and data interpretation is available through the fact that NEOTωIST is aimed at a known object. This object is the Itokawa asteroid (shown in Figure 2-1:), which has been thoroughly characterised by the Japanese Hayabusa mission. A further advantage of Itokawa is its elongated shape which makes a clean off-center strike easier, and causes clearly observable brightness fluctuations as it rotates.

Information about the impact geometry is substantially improved over what is known a-priori by observations from small sub-spacecraft which are deployed from the KI prior to impact. These sub-spacecraft also observe impact dynamics relevant to model validation (O5). The

resulting mission architecture is illustrated in

Figure 2-2: Overview of mission concept

The baseline architecture consists of:

The Kinetic Impactor, which performs all functions associated with the interplanetary transfer, performs terminal guidance navigation and control to the target, and acts as a carrier for the sub-spacecraft.

A Flyby Module (FBM) which is released onto a safe flyby trajectory prior to impact. The FBM performs observations of the impact from different positions along a trajectory roughly parallel to the impactor trajectory, including a view from 90° with respect to the impact velocity vector. This perspective yields more information about the geometry and dynamics of the ejecta cloud than for instance a view along the impact trajectory. After deployment, the FBM also functions as a data buffering and retransmission node for the other vehicles of the constellation.

Chasers (one or two vehicles) which are released from the Impactor and follow it along its trajectory, to be destroyed either by impacting the NEO or debris from the impact. This vantage point potentially provides an additional view of the ejecta cloud along its central axis. However, the unique observation opportunity that the Chasers offer is that of observing the impact crater, which the FBM cannot do because of obscuration of the crater by ejecta at the time when it may be geometrically possible. To have a chance of observing the impact crater, the Chaser must follow the Impactor with a sufficient delay (10s of seconds) to allow dispersion of most of the ejecta for a clear view on the crater. Crater characterisation is useful because it can constrain the



assumptions about total volume of ejected material and provide information about impact site properties.

Table 2-3: Information return by Mission Option & means of observation

Information MO-1 MO-2 MO-3 (Baseline)

Itokawa mass properties

Hayabusa mission Hayabusa mission Hayabusa mission

Spin rate change

Earth-based observation



Impact location

Impactor Nav-camera or targeting accuracy knowledge

Impactor Nav-camera & FBM observation

Impactor Nav-camera & FBM observation & crater imaging by Chasers

Ejecta moment-um vector direction

No observation; modelling based on asteroid geometry

FBM observation of ejecta geometry

FBM observation of ejecta geometry; pot. observation by Chasers

Ejecta mass No observation FBM observation; optical density of ejecta cloud

FBM observation; crater volume estimate from Chaser observation

Impact physics

No observation Ejecta geometry & expansion rate, density distribution; radiometry of blast

Ejecta geometry & expansion rate, density distribution from two vantage points and times; radiometry of blast

The baseline mission scenario consists of all elements described above (Mission Option 3). However, different implementation options with different levels of information quality are possible and can be chosen to match budget availability (C2). Other possibilities include the Impactor + Flyby Module (Mission Option 2) or only the Impactor (Mission Option 1). Table 2-3 shows what information is available and how it is obtained for the different Mission Options (MO). This provides an idea of what the descoping options mean in terms of mission return.

2.3 Sub-spacecraft deployment and observation strategy The primary observation target of the FBM during the flyby is the ejecta cone that develops after the impact. The high velocities and the observation geometries makes this a highly dynamic and challenging imaging task. A suitable image tracking strategy has to be selected that is able to cope with errors in the flyby geometry and with the uncertainties related to the ejecta cone (velocity and geometry).

Most of the operations take place in the last 64 hours of the mission (16h for terminal navigation + 48h for post-impact data downlink), with an observation core campaign that lasts only 10 seconds during the most critical part of the flyby. All these aspects regarding the FBM, its flyby geometry and interactions with other spacecraft are treated in the following section.

2.3.1 Flyby geometry

The flyby geometry, shown below, can be simplified by a right triangle with vertices the impact point (A), the point at offset distance from the impact point (B) and the separation point (C). In a first order approximation it has been assumed that the normal of the surface at impact point has



the same direction of the triangle side AC, whose length defines the release distance dr. This also means that the angle (∢BAC) is by definition assumed as 90°. The main parameters that drive the flyby geometry are the offset distance (doff) represented by the triangle side AB and the release distance.

Figure 2-3: Illustration of flyby geometry

To design the FBM subsystems a strategy has been developed that involves the development of a numerical model of the flyby event alongside with a simplified model of the ejecta cone. This has been performed in order allow preliminary design of the key spacecraft subsystems, which are highly dependent on the geometry chosen.

The ejecta cone is modelled in a first order approximation assuming linear expansion behaviour. The uncertainties in the cone geometry, that are also subjects of the observations to be made, appear in the half cone angle δ and in the characteristic velocity of the expelled material Vchar, which in turn depends on the impact velocity. To allow the FBM to sense the ejecta cone from different points of view in order to determine the geometry of ejecta cone it is important that the point of view vary fast enough in relation to the ejecta dynamics. For assessing this, the ejecta has been considered visible from a brightness point of view for a period of 10 s after the impact. The matching of the flyby geometry and the ejecta dynamics is essentially driven by the choice of the passage time (tpassage), which is the time interval between the impact and the passage of the FBM at point B. On the other hand the offset distance drives how fast the point of view changes during the observation phase.

To find appropriate values for these parameters an extensive test campaign has been performed and a good set of doff and tpassage has been chosen such as to fulfil the described considerations.

A reference scenario has been selected. Table 2-4: sums up the main parameters of the flyby numerical model of this scenario.

Table 2-4: Flyby model parameters

Parameter Value

doff 69.5 km

dr 28836 km

VIM 8.01 km/s

VFBM 8.0056 km/s

tpassage 2 s

ΔVN 4.45 m/s

ΔVS 19.29 m/s

δ 40÷50 °

Vchar 0.01÷0.1 VIM

tobs (core) 10 s

�̂�𝐴



Parameter Value

tobs (extended) 10 s

FOV 8.35°

# pixel 2048 x 2048

*N=Along track, S=Cross track (with respect to an impactor-centred reference system) To link together the flyby event, ejecta cone and camera parameters it is necessary to define a pivoting point or aiming point the camera should follow during the observation phase. This is a crucial step since the definition of such point influences many key subsystem parameters. In the previous phases of the project this point has been chosen to be the impact area on the asteroid surface. However, this definition does not take into consideration the ejecta dynamics and may end up sizing a bigger field of view (FOV) than required. For this reason the pivoting point has been defined in a more dynamic fashion as that point that is halfway between the top of the ejecta cloud and the base, on the cone edge line that is closer to the FBM during the flyby. The objective is a well centred view of the ejecta cone. Note that this definition of pivoting point proves to work well for the spacecraft design, but may not represent the absolute optimal solution. Its optimization is left for further study in future phases of the project.

To illustrate the flyby observation view, Figure 2-4 shows the ejecta cloud as seen from the spacecraft in 3 different moments of the core observation phase. The green area represents the half of the ejecta cone that is closer to the FBM while the red part is the other half that is further away. The white object is the target asteroid Itokawa. It is possible to see how the choice of the flyby geometry along with a proper timing of the events influences the point of view of the observation campaign.



Figure 2-4: Ejecta cloud and Itokawa seen from FBM at 0.8s, 2s and 10s after impact

In the reference scenario it is possible to observe the cone from different perspectives with a maximum viewing angle difference of about 58°, without excessive stresses on the turning manoeuvre or on the possible resolution as is shown later on.

It is worth pointing out that the instrument FOV has been chosen in relation to the ejecta cone in such a way that the ejecta cone is always in the FOV during the whole observation phase.

The observation phase has been divided into two sections, a core section that last from 0 s to 10 s (where 0s represent the impact time) and an extended section that is divided in 5 seconds before impact and 5 seconds after end of the core phase.

Figure 2-5 shows the changes of the minimum FOV to use depending on the offset distance chosen and on the uncertainties of ejecta behaviour. The lower family of curves represents the FOV for a characteristic velocity (ejecta expansion velocity) equal to 1% of the impact velocity while the upper one represents the case where the characteristic velocity is 10% of the impact velocity. It is visible that uncertainties on Vchar have a strong influence in the minimum FOV while uncertainties on δ (the ejecta half cone angle) have a smaller but not negligible influence. As it is possible to see from, increasing the offset distance has the effect of decreasing the required FOV and thus increases the size of the optics to use. Note however that the minimum required FOV it is driven by different factors depending on whether a close or far flyby is considered. In the first case, the minimum FOV is driven by the large changes of observation range (distance FBM to target) over the course of the observation phase. In the second case, the minimum FOV is mainly determined by the rapid expansion of the ejecta cone seen from an almost fixed distance.

Figure 2-5: Minimum FOV required during observation phase as a function of offset distance, tested for

different ejecta dynamics parameters



As previously pointed out, a strong dependency on the ejecta cone parameter uncertainties has been generally observed for all the parameters of the numerical model. The worst case scenario in terms of ejecta uncertainties has been considered in the reference scenario. The result is that a FOV of at least 8.35° is necessary to observe the ejecta cone, when assuming it has a wide angle and fast expansion velocity.

Another fundamental result of the numerical model that drives the design of the spacecraft is the manoeuvre required to rotate the camera FOV around the pivoting point. Defining the azimuth angle as the angle formed by the vector from the FBM to the pivoting point and the trajectory line it is possible to investigate the slewing performance required for flyby observation. Figure 2-6 shows the azimuth profile and azimuth rate needed to follow implement the reference observation scenario. The profile is shown only for a short time interval after impact (that occurs at 0 s) to highlight the magnitude of the maximum azimuth rate. This outcome from the numerical model translates into a critical system requirement to take into account for the spacecraft design.

Figure 2-6: Azimuth profile and azimuth rate

Figure 2-7 shows the change of a pixel size projected from the FBM to the pivoting point distance for flyby scenarios with different offset distances. It is possible to see how generally performing the flyby further from the impact area has the disadvantage of increasing the minimum pixel size but has at the same time the non-negligible advantage to decrease the variation of the resolution over a longer period of time. On the other hand a closer flyby would result in a better minimum sample distance but only for a limited amount of time.

Figure 2-7: Projected pixel size at pivoting point distance during the flyby for different offset distances and δ.



Based on the outlined results it is possible a reference scenario with and minimum flyby distance of 69.5 km is considered a good compromise. This scenario allows a relatively large change of viewing perspective (about 58°) during the 10s core observation phase, with a good level of pixel size over time, a reasonable azimuth peak rate (~ 6.9°/s) and an adequate FOV (8.35°) and size of the optics.

2.3.2 Tracking strategy

A tracking strategy for the flyby scenario requires certain flexibility to cope with the possible uncertainties and errors given by the real flyby geometry. In this sense the possibility to feed the system directly with a fixed azimuth profile as reference command to follow in an open loop fashion is not robust, or would require a very large field of view at the expense of resolution.

A more robust approach proposed for this mission is a tracking strategy consisting of two phases, a geometry determination phase / navigation phase (T1) and an open loop tracking phase (T2).

2.3.2.1 Geometry determination phase (T1)

During this phase observations about the real flyby geometry are collected, as the FBM approaches the target. Possible observations can be divided into two groups, the ones collected by the FBM and the ones taken by the KI and exchanged with the FBM.

For the latter group it is possible to make use of the fact that the KI position error relative to the target constantly decreases between FBM separation time and the precision impact. It is therefore possible to use the communication link between the two spacecraft to communicate to the FBM a history of the correction manoeuvres performed by the Impactor after separation. If the uncertainties of the ΔV at separation are well known it is possible for the FBM to use this information to solve the flyby geometry.

In principle the communication link between the two spacecraft itself could be used as additional source of information that gives relative distance and velocity information between the two (not currently baselined).

Another possibility is to use the on-board instrumentation of the FBM to perform visual navigation from the time when the illuminated asteroid becomes visible to the payload camera (possible at subpixel apparent size). Assuming that after separation the FBM attitude puts the asteroid target within the field of view, a constant inertial orientation makes it possible to see the asteroid center of brightness moving around the sensor with a measurable velocity. This apparent motion provides information about the true approach geometry. As distance from the asteroid decreases the target asteroid will start to be imaged from tens to hundreds of pixels. At this point a more sophisticated algorithm such as a template matching technique could be used, comparing the real illuminated shape of the asteroid with the one from a catalogue. Information about the asteroid-pixel velocity, apparent size and shape as viewed in the sensor can be therefore used to solve the real flyby geometry. Whether only centre of brightness navigation or more complex approaches taking onto account the asteroid geometry are required is still the subject of further analysis.

The techniques cited above can be used in combination or in redundancy depending on the performances and computational efforts required. The current baseline is trajectory reconstruction based on visual observations during approach.

It is important to note that all of the techniques discussed above are not considered feasible for the phases of the flyby after the impact because of the loss of the KI and because of the unknown visual signature of the impact and ejecta cone. This may temporarily "blind" the sensors or require additional complexity in the algorithm design, compromising the turning manoeuvre robustness during the most critical part of the flyby. For these reasons we choose to make use of an open loop tracking strategy for the short observation phase lasting from immediately before the impact until the end of the observation.



2.3.2.2 4.2.2 Open loop phase (T2)

It has been explained how during phase T1 observations from different sources can be collected. The idea is to make use of these observations to best fit a numerical flyby model in order to determine the real flyby geometry. Once the main flyby parameters such as offset distance, passage time and release distance are determined it will be possible to determine the reference command that allows the sensor FOV to perform observation rotating around the predetermined pivot point.

2.3.3 Operations timeline

In the following figure a possible reference timeline involving all the spacecraft of the mission has been designed. T0 is the impact time, which has been set at 0s.



Figure 2-8: Flyby reference mission timeline

2.4 Analysis of flyby navigation

2.4.1 Position estimation (w/o camera noise)

Spacecraft stack

Start corrective maneuvers

T0-16h

T0-1.2h

T0-1h Separation

T0-55m

FBM IM CH

End corrective manoeuvres

T0-2m

T0-5m Start geometry reconstruction

T0-5s Start extended observation phase

T0

T0+2s Passage at B

T0+30s

T0+10s

T0+15s

T0+20m

T0+30m Start downlink

End downlink. End nominal mission phase

T0+48h

Attitude Maneuver

Start T1 phase

Start T2 phase

Start core observation phase

End core observation phase

End extended observation phase

Communication attitude maneuver



The figure below shows the geometry of the flyby problem

Figure 2-9: FlyBy geometry

the distance estimation Ri is done using two consecutive camera measurements αi and αi-1.

Assuming:

Time interval Last 300 seconds

Frequency of measurements 1Hz (every 1 second)

Noise of camera σcamera 0

the time profile of difference between the real trajectory and estimated one is showed in Figure 2-10.

As we can see, the numerical error generates a line with a noise. There is a mitigation of this error if we consider:

- a lower frequency of measurements (for example one camera measurement every 43 seconds) but using even two consecutive camera meas. ;

- or, combining two not consecutive measurements. In the simulation, assuming a measurement frequency of 1 Hz for all time of mission (7228 seconds, around 2 hours), we have 7228 measurements. For the first distance estimation, done at 6928 s (300 seconds before the end of mission), we assumed to take 1st and 6928th camera measurements. For the next distance estimation at 6929 s, I toke 2nd and 6929th measurements; for the next one, 3rd and 6930th measurements and so on.

Ri

Itokawa

i

i-1

αi

αi-1

END Ri-1

FBM trajectory



Both these approaches have the same purpose: considering a bigger “triangle” (in Figure 2-9 formed by i, i-1 and Itokawa).

Figure 2-10: Time profile of difference between real trajectory and estimated one, Freq. meas. 1 Hz (no

camera noise)

Figure 2-11: Time profile of difference between real trajectory and estimated one, Freq. meas. 0.023 Hz (one every 43 sec) (no camera noise)



Figure 2-12: Time profile of difference between real trajectory and estimated one, considering not two consecutive camera measurements (no camera noise)

Looking in details the figures:

- Figure 2-10, good order of error ≈ 10-9 km, but very noisy; - Figure 2-11, best order of error ≈ 10-10 km, just a bit noisy; - Figure 2-12, “worst” case, because the order of error is ≈ 10-7 km, but the signal is super

clean.

As we can see, considering a perfect camera (w/o noise) these three approaches have different pros and cons. To choose, the best one it is necessary to check the performance when you add the camera noise, to see also the influence of the numerical error on distance estimation.

2.4.2 Position estimation (with noise)

We started considering the first case, namely:

Time interval Last 300 seconds

Frequency of measurements 1Hz (every 1 second)

Distance estimation based on

Two consecutive camera measurements

We assumed to introduce a little noise in camera measurements: 1 arcsec. The result is showed in Figure 2-13. Obviously, the performance is poor and there is a big error. To mitigate it and before using a Kalman Filter, we designed a simple low pass filter, to see effectively if we have an improvement.

For definition, the low pass filter passes signals with a frequency lower than a certain cutoff frequency and attenuates signals with frequencies higher than the cutoff frequency. This is determined analyzing the Power Spectal Density (PSD) of camera measurements (Figure 2-14). PSD shows the strength of the variations (energy) as a function of frequency. In other words, it shows at which frequencies variations are strong and at which frequencies variations are weak. Therefore, in our case the cutoff frequency is that one where the PSD is still smooth. In this case that is 10-2 Hz.



Figure 2-13: PSD of camera measurements, Freq. meas. 1 Hz (1 arcsec noise)

Figure 2-14: Time profile of difference between real trajectory and estimated one, Freq. meas. 1 Hz (1 arcsec)

In Figure 2-13 we can see that the Low Pass Filter is able to mitigate the noise but at the end there is a sort of “shift”. Basically the orange signal converges to a value that it’s not zero. This happens because the Low Pass Filter mitigate only the frequency, doesn’t care about the dynamics, it’s normal for definition.

Therefore, with Low Pass filter shows that it’s possible to mitigate the noise of measurements, but at the same time, we need a more powerful filter: Kalman Filter.



Setting an initial position uncertainty of 100 km and an initial velocity uncertainty of 0.1 km, the results of Kalman Filter are the following (Figure 2-15, Figure 2-16, Figure 2-17):

Figure 2-15: Time profile of difference between real trajectory and estimated one, Freq. meas. 1 Hz (1 arcsec)

As we hoped the Kalman Filter is able to mitigate a lot the difference between the real trajectory and the estimated one. Most important the behavior of different quantities (position and range) became very smooth. That is important because in this way is possible to develop a more consistent rotation law of the camera.



Figure 2-16: Time profile of distance from asteroid

Figure 2-17: Time profile of distance from asteroid (zoom)

2.5 Communications concept A core feature of the NEOTωIST concept is its communications architecture. During the cruise of the spacecraft stack all TT&C is performed via a link to the Impactor. In the terminal phase shortly before and after impact, multiple commination links connect the different sub-spacecraft and Earth, as summarized in the table below. The FBM functions as central data buffering and



communications node for the constellation, with most of the payload data being downlinked to Earth from the FBM after the impact and flyby.

Table 2-5: Overview communications links

Link Type Purpose Impactor <-> Earth

X-band, low rate

Telemetry and select compressed images from navigation camera for redundancy

Impactor -> FBM

X-band, high rate

Telemetry & manoeuvre history for relative trajectory estimation between KI and FBM; KI Nav-camera images for later downlink to Earth

Chaser ->FBM

X-band, very high rate

Transmission of chaser imagery in real time (transmission before Chaser destruction)

FBM <-> Earth

X-band, medium rate

Delayed downlink of telemetry and observation data from the terminal phase, for all sub-spacecraft

A summary of the data budget and link speeds (inter-satellite and payload data dump to Earth) is given in Table 2-6. Some of the inter-satellite (ISL) link speeds are challenging. Feasibility is supported by the fact that inter-satellite communications need to be maintained only for short periods, which enables the use of very high Tx power (energy availability and thermal design).

Table 2-6: Data budgets and link speeds

2.6 Design of Impactor and spacecraft stack The KI design is envisioned as an updated version of the Impactor design for the Don Quijote (DQ) mission. The concept of the DQ impactor is shown in in the figure below. The spacecraft consists of a large propulsion module which was flown for the Lisa Pathfinder (LPF) Mission, and a small Mission Module which is mission-specific and performs other functions required for the mission (interplanetary transfer and terminal impact GNC).

FBM* KI** CH***Spacecraft Data Frames/s 15 (5) 1 1# Pixel 2048 x 2048 1024 x 1024 2048 x 2048Obs time [s] 10 (+10) 40 30Payload data[Mbit] 7200 250 720Telemetry [Mbit] 200 1 3.6Data [Mbit] 7400 251 723.6Inter-spacecraft linksMax range [km] / 72 235Band X X XComm. Window [s] / 20 + 40 30Link data rate [Mbit/s] 50 6.25 24.1Antenna type HGA & MGA MGA MGA

Downlink performance

Range [AU]Data rate [Kbps]Data[Mbit]Downlink time [h]

150

8374.646.53



Figure 2-18: Impactor concept, based on Don Quijote design

The only modification to the LPF propulsion module is the addition of manoeuvring thrusters needed for terminal GNC control authority.

Main updates of the Mission Module consist of tailoring power and communications systems to the mission geometry. The communications system can be simplified since only low-rate transmission of TT&C and some compressed images are required between the KI and Earth. Avionics, in particular power and data handling, are updated to save mass and power. Table 2-7 shows a top-level mass breakdown of the spacecraft stack.

Table 2-7: .Top-level spacecraft stack mass budget

The masses shown contain maturity margin between 5% and 20%, as well as 20% system margin. For some mission opportunities the combined mass of Mission Module, FBM, and Chaser is marginal. The VEGA-C launcher upgrade will add margin in all cases.

2.7 Design of Flyby Module The Flyby Module (FBM) functions as the main imaging platform for the impact event, and as a data buffering and communications node for the satellite constellation. Its design is considered a key challenge of the proposed concept.

2.7.1 7.1 Design drivers for FBM

The main design drivers for the FBM are:

Table 8. Flyby module design drivers and implications

Driver Design implications

High quality imaging of impact event during high velocity pass

- Relatively powerfully optics - Fast actuation of image tracking - Precise and flexible image tracking GNC

Dry [kg] Wet [kg]Fly-by Module 104.1 104.6Chaser 25.7 26.2Impactor Mission Module 263 263Propulsion module (incl. resid.) 280 1420Stack 673 1814Stack at launch incl. adapter 1929Delivered mass w/o prop. mod. 394Min. impacting mass 543

Vehicle stack mass synopsis




- Precision inertial stabilisation during imaging Data buffering & transmission to Earth for entire constellation

- Robust high-rate ISL links - Powerful Tx capability for Earth downlink - Attitude control authority for duration required to dump

payload data Size and mass constraints imposed by overall mission budget

- Small satellite design approach (integration) - Low mass equipment selection - Critical approach to redundancy

Mission cost constraints - See above - Use of existing equipment where possible

Deep space environment - Radiation tolerance for some attractive LEO equipment to be checked

Next do these considerations the design of the FBM is influenced by the decision to not pursue a solution which is highly integrated with the primary mission vehicle, here the KI. This leads to a certain amount of redundancy between both vehicles, for instance in communications or power generation. The rationale for this is mainly programmatic with the following main reasons:

- The ad-on character of the FBM allows flexibility in the choice of whether to implement it or not (flexibility wrt. budget constraints)

- Vehicle interfaces and design are simper (no cross feeding of power, less data & control interfaces, less configuration constraints).

- The developed solution can be used as a basis for add on capabilities on other mission scenarios (e.g. very low flybys of planetary objects for imaging or sampling) -> Multi-mission platform

2.7.2 7.2 Flyby module payload

The main payload of the FBM consists of a panchromatic camera with a 2048 x 2048 sensor. Since the peak azimuth rate of the reference scenario is high (6.9 °/s), the use of a one-axis rotating pointing mirror instead of an attitude manoeuvre is considered a more appealing solution. Figure 2-19 illustrates the proposed solution for the FBM field of view tracking.

Figure 2-19: Payload subsystem with external pointing mirror

The use of a pointing mirror has several advantages over an attitude manoeuvre. Less mass is rotated, which means that less power and actuator torque is needed. An vehicle attitude manoeuvre would require very large reaction wheels or control moment gyros (CMGs), both solutions that were found to be less attractive than the proposed solution in terms of power and



mass. Alternatively using a propulsive approach is problematic due to induced vibrations and the non-constant rate profile during imaging. Given these constraints it is easier, stable and cheaper to design a system where only an optical component of the payload is put in rotation and not the whole spacecraft. A solution with an external circular rotating mirror is therefore proposed for the FBM camera.

Next to the camera the FBM features a Radiometer which samples the brightness of the impact at high frequency (kHz). Since it is non-imaging, the optics and detector are compact and simple. Target pointing is integrated with the main camera, whereby pointing requirements are much less stringent. High-frequency radiometric measurements provide information about dynamics of the impact event during the first milliseconds.

2.7.3 7.3 Outline FBM platform and operations

The table below gives an overview of the FBM specifications and equipment type. Table 2-9: Flyby module specifications

Item Description

Mass Dry/ Wet: 104.1 / 104.6 kg

Power DC: 100 W

S/A size: 0.75 m²

Battery: 150 Wh (usable)

Dimensions Bus structure: 100 x 100 x 25 cm³

Payload Medium angle camera

Field of view: 8°

Aperture: ~ 9 cm

Detector: 2048 x 2048 pixel

Max. resolution at 70 km: 5m

Field of view pointing: pointing mirror

Image targeting: based on visual navigation using payload camera

Non-imaging radiometer

Power and data handling

Integrated power and data handling (e.g. PROBA-NEXT), >1 Gbit mass memory

Comms. Earth-link: X-band, 1m HGA, 20 W RF, 50 kb/s

Impactor link: X-band, MGA, Rx only, < 10 Mbit/s

Chaser link: X-band, MGA, Rx only, ~ 20 Mbit/s

AOCS 3-axis stabilised

2 x STR (DTU Micro ASC)

2x Coarse sun sensors

3 x Micro-wheels, 0.42 Nms (SSTL)

IMU (DTU, int. with STR electronics)

RCS: 4 x cold gas thrusters for momentum management (0.5 kg of N2)

Thermal Heaters & radiators The figures below show the configuration concept of the FBM.



Figure 2-20: Flyby module configuration

Figure 2-21: Flyby module equipment configuration

The FBM configuration is roughly a box shape (dimensions 100 x 100 x 25 cm³), with the mounting and orientation of the equipment driven by the communications and imaging geometry. This geometry is illustrated in Figure 2-22. Note that the angle between high gain antenna and solar array must be adjusted according to the specific mission opportunity. The FBM design allows this via and antenna mounting that can be adjusted accordingly before launch.



Figure 2-22: FBM communications and imaging geometry during flyby and data downlink

The FBM attitude is inertially fixed using the reaction wheels and star trackers during the flyby. Dynamic field of view pointing in azimuth is performed by a pointing mirror as described above. Optimal fixed elevation is acquired by spacecraft attitude adjustment. The flyby attitude is adjusted after release from the Impactor in two steps:

Coarse orientation: Based on known approach trajectory, attitude is adjusted such that target NEO is in field of view.

Fine adjustment: based on visual tracking of target NEO during approach attitude is adjusted to place target object at 0° elevation (centre of field of view) during flyby.

After the flyby, the FBM is oriented for optimal power generation and data dumping to Earth. The cold gas propulsion system is intended to provide momentum management capability during this phase.

2.8 Design of Chaser spacecraft The Chaser functions as a secondary imaging platform with the main goal of imaging the impact crater after the ejecta cloud has dispersed sufficiently to be transparent. In addition it may perform long range ejecta observations from a secondary vantage point. The significant temporal delay required to allow ejecta dispersion requires a relatively large along-track separation between Chaser and the Impactor, which in turn requires a relatively large delta-V after Chaser separation. Since the Chaser follows the Impactor without lateral offset, the default scenario is its destruction by collision with either the target NEO or ejecta debris.

2.8.1 8.1 Design drivers for Chaser

The main design drivers for the Chaser are:

Table 2-10: Chaser design drivers and implications


Uncertainty in imaging geometry

& small target

- Wide field of view Relatively high resolution (large number of pixels)

Real-time transmission of imaging data before

- Robust very high-rate inter-satellite link




vehicle destruction

Large delta-V for separation from Impactor

- High-Isp propulsion solution

Mission cost and mass constraints

- Small satellite design approach - Critical approach to redundancy - Use of existing equipment where possible - Simple AOCS approach

Deep space environment - Radiation tolerance for some attractive LEO equipment to be checked

2.8.2 8.4 Outline Chaser design

The table below gives an overview of the Chaser specifications and equipment type. Table 2-11: Chaser specifications

Item Description

Mass Dry/Wet: 25.7 / 26.2 kg

Power DC: 40 W

Battery: 60 Wh (usable)

Dimensions 20 x 30 x 10 cm³

Payload Medium angle camera

Field of view: 16°

Aperture: still under analysis

Detector: 2048 x 2048 pixel

Max. resolution: 4m @ 30 km, 2m @ 15 km

No active target tracking

Power & data Cube-sat/ small sat equipment (details still to be selected)

Comms. Link to FBM: X-band, MGA, TX only, ~ 20 Mbit/s

AOCS Stabilization along velocity direction with single uncontrolled momentum wheel, spun up before ejection from Impactor

Propulsion Hydrazine, single thruster in anti-velocity direction, stabilization with momentum wheel, ΔV capability 70 m/s

Thermal Heaters & radiators, non-stationary design for terminal phase possible

The figure below shows a preliminary sketch of a possible internal configuration of the Chaser and the relative position of the FBM and imaging target. A tentative shape goal is a 6U-Cubesat envelope.



Figure 2-23: Chaser configuration sketch

3 Impactor Guidance Navigation and Control Extensive impactor GNC testing was performed for the NEOTωIST case. The results are summarized in [RD1].



4 Observation and measurements

4.1 Target Asteroid The lack of a reconnaissance spacecraft in this relatively inexpensive alternative mission would reduce the scientific return of the mission. However, using a previously visited asteroid as a target would partially compensate for this, since some scientific context for the mission would already be available. Possible targets in this case could include (25143) Itokawa, (101955) Bennu, and (162173) Ryugu. Of these 3 asteroids only Itokawa has been visited to date, when the JAXA Hayabusa mission observed Itokawa from a distance of the surface to tens of km away throughout many months in 2005. The information collected then, can be used to replace some of the functions of the observer spacecraft in a two spacecraft mission.

Itokawa is a 535 x 294 x 209 m large stony type (IV) asteroid (Fujiwara et al., 2006) with a mass of 3.51 x 1010 kg. The elongated shape combined with an almost perpendicular to the ecliptic spin axis makes it possible to measure the rotational period of 12.132370 hrs (Scheeres et al., 2007) to a high precision, and therefore also changes to it.

4.2 Observation principle and analysis The observation principle is discussed and analysed in detail in [RD3].

4.3 Observability of momentum enhancement factor (beta) A detailed analysis was performed on the feasibility and achievable accuracy of determining the momentum enhancement factor (beta factor) with the NEOTωIST mission concept. The results are summarized in [RD4].

4.4 Observing the ejecta in the visual The optimal illumination angle to observe the ejecta cloud will depend on the size of the particles in the cloud, as the light scattering on particles depends on the ratio of the particle radius, r, to the wavelength, λ. Using the formula a = 2 π r/λ then for a > 50 geometric scattering will produce backwards scattering primarily. In the case of 0.1 ≤ a ≤ 50 Mie scattering will dominate, creating forward scattering, and for very small particles with a << 1, it is primarily Rayleigh scattering, where the light is scattered equally forward and backward. (Andrews, 2010).

To have primarily backward scattering in the visual part of the electromagnetic spectrum (0.4-0.7 μm) then the particle radius of the clouds needs to be larger than 3-5.5 μm. For it to be forward scattering the radius needs to be smaller than that, but larger than 0.006-0.011 μm.

To calculate the apparent magnitude of an ejecta cloud, the flux from all particles are calculated and added together and the total flux is converted back to apparent magnitude, m:

𝑚 = −2.5 𝜅�𝑙𝑙𝑙10(10−0.4 𝑚1 + 10−0.4 𝑚2+. . . +10−0.4 𝑚𝑛)

where κ is an opacity factor.

4.5 Observing the impact in the infrared A thermophysical model of Itokawa was created and run for the 5 impact times in the years from 2022-2034. Simulated surface temperatures are used to calculate the thermal infrared signal from the asteroid for estimation of the signal to noise of observing the impact event.



Based on (Demura et al., 2006) and (Miyamoto et al., 2007) an overall idea of the location of the smooth terrains on Itokawa is gained. Furthermore, close-up AMICA images (http://darts.isas.jaxa.jp/planet/project/hayabusa/amica) are used to outline in detail the edges between the smooth and rough terrains on the Gaskell 3D shape model in Meshlab. The resulting terrain map was used to update the asteroid surface thermal model by I. Pelivan (Pelivan et al., 2017) to differentiate between areas with different thermal inertias. We have assigned the regions colored in black a thermal inertia of 500 J m-2 K-1 s-1/2 and all other terrain a thermal inertia of 900 J m-2 K-1 s-1/2, see Müller et al, 2014). The mission’s target point lies in the latter region. The terrain map is loaded in conjunction with the shape model to assign each shape facet the associated thermal inertia value.

The center facet of one of the target eclipse(s) shows a diurnal temperature variation peaking close to 335K with a minimum of around 210 K for the possible target date of Febuary 12, 2027 (see Drube et al for all possible target dates). The diurnal surface temperature variation for this target date is shown in the figure below.

Figure 4-1: Diurnal temperature variation for a target facet of the 50K version of the Gaskell shape model for Febuary 12, 2027.

Surface temperatures have been calculated for all target dates and all shape facets (see figure below).

Figure 4-2: Surface temperature distribution on Itokawa 9.27 hours into the rotation, i.e. at start of day for the target facet of above Figure 4-1. Red facets have maximum temperatures above 300K, dark green represent low temperatures of ~200K, lighter colored facets have intermediate temperatures.

http://darts.isas.jaxa.jp/planet/project/hayabusa/amica



5 Baseline mission trajectories The analysis for potential mission opportunities has been performed assuming a launch on VEGA, which is selected as a low-cost European launch option. It places stringent mass constraints on the mission. As shown in the table below, good Mission opportunities exist at regular intervals of approximately 3 years. Transfer duration is on the order of 2-3 years in most cases, and communications distances at time of impact are manageable at below 1 AU, sometimes significantly less.

Table 5-1: Mission opportunities and parameters

Arrival year 2024 2027 2030 2033 2036 Flight time (days) 581 942 589 95 503

Earth V∞ (km/s) 2.07 2.39 2.77 2.18 1.62

Arrival V∞ (km/s) 8.01 8.48 8.89 8.98 8.93

Payload mass * (w/o LPF, kg) 424 401 372 416 447

Sun-Itokawa-impactor angle at impact (deg) 30 24 21 12 13

Earth-Itokawa distance at impact (AU) 0.95 0.68 0.39 0.09 0.25

Earth-Itokawa-impactor angle at impact (deg) 90 90 90 74 42

* Note that the payload mass refers to the mass delivered excluding the propulsion module, i.e. mass available for Impactor Mission Module, FBM, & Chaser

The 2030 mission opportunity is marginal in terms of delivered payload mass. The mass situation for this and all other mission opportunities will be somewhat relaxed once the upgraded VEGA-C launcher goes into service. The mission is also compatible with other larger launch vehicles.

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